Rocket assembly ablative materials formed from solvent-spun cellulosic precursors, and method of insulating or thermally protecting a rocket assembly with the same

ABSTRACT

A rocket motor assembly is insulated or thermally protected with a rocket motor ablative material formed from a prepreg. The prepreg contains at least an impregnating resin matrix and, as a precursor prior to carbonization, either carded and yarn-spun solvent-spun staple cellulosic fibers, solvent-spun cellulosic filaments, or a combination thereof. When patterned and carbonized, the rocket motor ablative material can be lined or otherwise placed into a rocket motor assembly, such as between the solid propellant and case, in the bulk area of the exit nozzle liner, or at susceptible portions of a re-entry vehicle, such as the nose cone.

RELATED U.S. APPLICATIONS

Priority is based on U.S. Provisional Application No. 60/124,670 filedon Mar. 16, 1999 and U.S. Provisional Application No. 60/097,117 filedon Aug. 19, 1998, the complete disclosures of which are incorporatedherein by reference.

Certain aspects of this invention were made under contract NAS 8-38100with the National Aeronautics and Space Administration.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates to rocket motor ablative materials, especiallyresin-filled carbon fiber and carbon/carbon ablative materials, and amethod of making the ablative materials. In particular this inventionrelates to carbon ablative materials having a reinforcement componentformed from, as a precursor prior to carbonization, carded and yarn-spunsolvent-spun cellulosic fibers and/or solvent-spun filaments. Thisinvention also relates to rocket motor assemblies including the carbonablative materials.

2. Description of the Related Art

It is generally accepted current industry practice to prepare insulationfor solid propellant rocket motors from a polymeric base compositeimportantly including a carbon cloth. The composite is generallycomposed of the carbon cloth as a woven reinforcement structureimpregnated with a suitable resin matrix. The resin matrix is commonly aphenolic resin, although other resin matrices can be used. For makingthe woven reinforcement structure, current industry practice is toselect a continuous filament of non-solvent-spun viscose rayon as aprecursor material (hereinafter referred to as continuous filamentviscose rayon). The continuous filament viscose rayon, which isespecially formulated for ablative applications, is woven, wound, orotherwise manipulated into its desired configuration and then carbonizedto form a carbon structure exhibiting superior ablation characteristicsand excellent physical properties and processability.

Continuous filament viscose rayon precursor has been established as astandard in the rocket motor industry for making carbon reinforcedstructures of carbon and carbon/carbon ablative materials due to itssuperior ablation characteristics, excellent physical and thermalproperties, and high processability. One of the excellent physicalproperties possessed by composites formed from continuous filamentviscose rayon precursor is a cured composite high warp strength of about144.8 MPa (about 21,000 lbs/in²) at ambient temperature (about 21° C. or70° F.), as measured subsequent to carbonization and impregnation of theprecursor. Warp strength reflects the tolerance of the filament toopposing forces acting along the warp (or longitudinal) filament axis.

However, a major drawback associated with the use of cured compositescomprising wrapped layers of continuous filament viscose rayon, such asfound within the bulk areas of much rocket nozzle insulation, is therelative low across-ply tensile strength possessed by the carbonizedcontinuous filament viscose rayon at operating temperatures experiencedwithin the bulk ablative material (as opposed to the exhaust gassurface) during firing of a rocket motor. Such firing temperatureswithin the bulk ablative material generally can rise to about 400° C.(or 750° F.). Specifically, cured composites comprising wrapped layersof carbonized continuous filament viscose rayon have across-ply tensilestrengths on the order of about 2.07 MPa (about 300 lbs/in²). Asreferred to herein, across-ply tensile strength is the amount of load,perpendicular to the filament axes, which two overlapping layers offilaments can withstand prior to slippage.

Another significant drawback associated with continuous filament viscoserayon that has recently drawn significant attention involves theavailability of this particular type of continuous filament. Over thepast few years, the only manufacturer producing sufficient quantities ofnon-solvent-spun continuous filament viscose rayon to meet industrydemands is North American Rayon Corp. (NARC) of Elizabethton, Tenn. Thecapability of the industry to produce ablative liners and other thermalinsulation based on continuous filament viscose rayon has beenjeopardized, however, due to the cessation of continuous filamentviscose fiber production by NARC. There is therefore a need in thisindustry, previously not satisfied, to find an effective alternatesource or a replacement candidate for the above-described standardthermal insulation formed from continuous filament viscose rayonprecursor.

The requirements that a replacement candidate must satisfy in order tobe acceptable and functionally effective are well known to be quitesevere due to the extreme conditions to which the insulation is exposed.These conditions not only include exceedingly high temperatures but alsosevere ablative effects from the hot particles (as well as gases) thattraverse and exit the rocket motor interior, or over the outer surfaceof re-entry vehicle insulators. Unless the insulation will withstandsuch conditions, catastrophic failure may (and has) occurred.

Accordingly, any replacement insulation should exhibit comparabletemperature resistant and ablation characteristics and rheological andphysical properties at least equivalent to those of continuous rayonviscose filament, yet should not otherwise significantly alter themanufacturing process employed for the production of the thermalinsulation. Additionally, due to the large and growing quantities ofsolid propellant rocket motor insulation required by the industry, anysuch replacement reinforcement precursor candidate should be abundantlyavailable now and into the foreseeable future.

An alternative carbon precursor that has been proposed for ablativeapplications is continuous filament polyacrylonitrile (PAN). PANcontinuous filaments disadvantageously possess higher densities thancellulosic materials (1.8 g/cm³ for PAN, compared to 1.48 g/cm³ forcellulosic filaments) and higher thermal conductivities than cellulosicmaterials. Thus, in order to provide a comparable insulation performanceto rayon filaments, rocket motor nozzle insulation or re-entry vehicleinsulation formed from PAN filament must have a greater thickness andweight than a comparable-performing insulation formed from cellulosicmaterials. The replacement material must meet the ablation limits forprotection of the casing (when used as an internal casing insulation)throughout the propellant burn without adding undue weight to the motor.

Accordingly, the search for a functionally satisfactory precursor formaking the reinforcement structure of a composite material requiresdiscovery and implementation of an extraordinarily complex combinationof characteristics. The criticality of the material selection is furtherdemonstrated by the severity and magnitude of the risk of failure. Mostinsulation is of necessity “man-rated” in the sense that a catastrophicfailure can result in the loss of human life—whether the rocket motor isused as a booster for launch of a rocket or is carried tacticallyunderneath the wing of an attack aircraft. The monetary cost of failurein satellite launches is well-publicized and can run into the hundredsof millions of dollars.

Therefore, one of the most difficult tasks in the solid propellantrocket motor industry is the development of a suitable, acceptableinsulation that will meet and pass a large number of test criteria tolead to its acceptability.

Furthermore, any replacement precursor should not be susceptible toobsolescence issues nor discontinuance in future supply thereof.

SUMMARY OF THE INVENTION

It is, therefore, an object of this invention to address a crucial needin the industry to reformulate the ablative liners and thermal liners ofrocket motors by finding a suitable replacement precursor for makingcarbon-based reinforcement structures. As referred to above, suitablereplacement means a precursor material that can be substituted forcontinuous filament viscose rayon without requiring significant amountsof modification to the impregnating resin composition, component design,and manufacturing process steps and, when carbonized, possess equal orsuperior properties, in particular overall strength, as the thosepossessed by the continuous filament viscose rayon standard.

In accordance with the principles of this invention, these and otherobjects of the invention are attained by the provision of a rocket motorablative material (e.g., an insulation liner or the like) formed from,as a precursor of the carbon reinforcement structure, yarn comprisingeither carded and yarn-spun solvent-spun cellulosic (e.g., rayon)fibers, solvent-spun cellulosic filaments, or a combination thereof. Theinventors discovered that solvent-spun cellulosic fibers and filamentsare capable of being processed, such as by spinning, into yarns which,upon patterning (e.g., weaving, in any weave style or winding) andsubsequent carbonization, can serve as a reinforcement of prepregs andcan be processed into an insulation liner under conditions comparable tothose of continuous filament viscose rayon.

The inventors also discovered that when carded and yarn-spunsolvent-spun staple cellulosic fibers possessing certain dimensionalcharacteristics are selected, the resulting yarn possesses excellentmechanical strength for rocket motor applications, yet does not releaseunacceptable levels of fiber fly—i.e., short, waste fibers—into the airin textile processing operations such as carding, yarn-spinning, andweaving. The former discovery, in particular, was especially surprisingbecause yarns prepared from cellulosic fibers were expected to possessand do possess significantly lower warp strengths than yarns producedfrom continuous filament viscose rayon. However, the inventors foundthat the lower warp strengths of the yarns prepared from cellulosicfibers are compensated for by the far superior across-ply tensilestrength that yarns prepared from cellulosic fibers exhibit overcontinuous filament viscose rayon.

This invention is also directed to a rocket motor assembly comprisingablative materials which comprise reinforcing structures formed from, asa precursor material prior to carbonization, yarn comprising eithercarded and yarn-spun solvent-spun cellulosic (or rayon) fibers,solvent-spun cellulosic filaments, or a combination thereof. Thisinvention is further directed to a process for making a rocket motorassembly comprising the ablative materials including nozzle and re-entryvehicle components.

Other objects, aspects and advantages of the invention will be apparentto those skilled in the art upon reading the specification and appendedclaims which, when taken in conjunction with the accompanying drawings,explain the principles of this invention.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings serve to elucidate the principles of thisinvention. In such drawings:

FIG. 1 is a schematic cross-sectional view depicting the insulation ofthis invention interposed between a rocket motor casing and solidpropellant;

FIG. 2 is a schematic cross-sectional view identifying some of theregions of a rocket motor assembly in which the insulation of thisinvention may be applied; and

FIG. 3 is a flow diagram of a process for making LYOCELL fibers.

DETAILED DESCRIPTION OF THE INVENTION

In accordance with the principles of this invention, the replacementprecursor material for preparing carbon reinforcement structures ofrocket motor ablative materials, including nozzle re-entry vehiclecomponents, is yarn comprising either carded and yarn-spun solvent-spuncellulosic (or rayon) fibers, solvent-spun cellulosic filaments, or acombination thereof. As referred to herein and understood in the art,carded means fibers subjected to a process or passed through a machinedesigned to promote the at least partial separation and at least partialalignment of fibers. Carding encompasses techniques used in theproduction of both fine and coarse yarns. As referred to herein andunderstood in the art, yarn-spun means a yarn formed a combination ofdrawing or drafting and twisting of prepared fibers. Yarn-spinning asreferred to herein is not intended to mean techniques consisting of theextrusion of continuous filaments, which techniques can be performedduring solvent-spinning. As referred to herein, staple fibers are fibershaving lengths suitable for yarn-spinning.

Various solvent-spun cellulosic fibers can be used in accordance withthis invention. A representative example is of a solvent-spun cellulosicfibers are LYOCELLcellulosic fibers made by spinning withN-methylmorpholene-N-oxide.

The solvent-spun cellulosic fibers preferably have average fiber lengthsin a range of from 38 mm to 225 mm, such as 100-150 mm. The solvent-spuncellulosic fibers, when processed into a yarn, and/or solvent spun intofilaments preferably have an average denier per filament (dpf) in arange of from 1.1 dpf to 3.0 dpf. One supplier of solvent-spuncellulosic fibers and filaments is Lenzing Fibers of Austria. Thesolvent-spun cellulosic materials provided from this commercial sourcegenerally have sodium and zinc levels of 90 ppm and 2 ppm, respectively,which are less than the 1300 ppm and 300 ppm of typical continuousfilament viscose rayon supplied by NARC. Solvent-spun cellulosic fibersmade by solvent-spinning with N-methylmorpholene-N-oxide are commonlyknown as LYOCELL.

The cellulosic fibers and filaments are preferably untreated, meaningthat they are free of any distinct metallic, metalloidic, or graphiticcoating, at least prior to (and preferably subsequent to)graphitization.

One of the advantageous features of this invention is that the yarncomprising the solvent-spun filaments and/or the carded and yarn-spunsolvent-spun cellulosic fibers may be substituted for conventionalcontinuous filament viscose rayon without significantly altering theablative material manufacturing process. The only substantial alterationin the manufacturing process, at least with respect to the processing ofreinforcement structures from solvent-spun cellulosic fibers, resides inthe differences between producing the yarn of this invention andproducing conventional continuous filament viscose rayon. Generally,continuous filament viscose rayon is produced by dissolving celluloseinto a viscose spinning solution, and extruding the solution into acoagulating medium where the polymer is cellulose and is regenerated asa continuous filament. On the other hand, the yarn used in an embodimentof the present invention is prepared from solvent-spun staple fibers,which are carded and yarn-spun by techniques well known in the industryinto a tight, compact yarn from the staple fibers. It is understood thatother processing techniques may also be used, such as combing and othersteps well known and practiced in the art. Preferably, the yarn-spinningstep is performed by either a worsted process or cotton-ring spinningprocess. The yarn-spinning process is advantageous to keep yarnhairiness to a minimum. By way of example, the yarn may have a weightcomparable to the weight of standard yarns presently used for carbonablative materials, i.e., about 1650 denier. This may be accomplishedwith staple fibers by producing a yarn that is approximately 4.8 Englishworsted count (Nw), and two-plying the yarn to obtain the 1650 denierconfiguration. Suitable amounts of twist attained by spinning can be,for example, 2-12 360° turns per inch, more preferably 10-12 360° turnsper inch.

The yarns are then subject to one or more patterning techniques,including, by way of example, weaving, winding, and plying, into adesired structure. The structure is then carbonized to form thereinforcement of the ablative material. In this regard, the structuringof the yarns into the desired configurations can be performed in thesame manner as that for conventional continuous filament viscose rayon.Carbonization can take place, by way of example, at a temperature of atleast 1250° C., preferably at least 1350° C. The carbonizedreinforcement structure is then impregnated with an acceptable resin,such as a phenolic resin. A representative phenolic resin is SC1008,available from Borden Chemical of Louisville, Ky.

The inventive ablative and insulation materials can be applied tovarious parts of a rocket assembly, preferably as multi-layeredstructures. For example, the ablative and insulation materials can beused as a chamber internal insulation liner, as shown in FIG. 1.Referring to FIG. 1, the insulation 10, when in a cured state, isdisposed on the interior surface of the rocket motor case 12. Typically,a liner 14 is interposed between the case 12 and the insulation 10. Theinsulation 10 and liner 14 serve to protect the case from the extremeconditions produced by the burning propellant 16. Methods for loading arocket motor case 12 with an insulation 10, liner 14, and propellant 16are known to those skilled in the art, and can be readily adapted withinthe skill of the art without undue experimentation to incorporate theinsulation of this invention. Liner compositions and methods forapplying liners into a rocket motor case are also well known in the art,as exemplified by U.S. Pat. No. 5,767,221, the complete disclosure ofwhich is incorporated herein by reference.

The ablative and insulation materials can also (or alternatively) beapplied along the flow path through which the combustion products pass,such as shown by the shaded area 20 of the exit nozzle shown in FIG. 2.

The ablative performance and mechanical properties of the carbon clothphenolic prepared from cellulosic fibers as a precursor to thereinforcement are comparable those of carbon cloth phenolics made fromthe aerospace-grade continuous filament viscose rayon in subscale testmotors. For example, although carbonized yarns formed from solvent-spuncellulosic fibers exhibit a slightly lower warp strength than yarnsformed from continuous filament viscose rayon (96.5 MPa (or 14,000lbs/in²) compared to 144.8 MPa (or 21,000 lbs/in²)), carbonized yarnsformed from solvent-spun cellulosic fibers have an across-ply strength(5.52 MPa or 800 lbs/in²) twice that of continuous filament viscoserayon (2.76 MPa or 400 lbs/in²) (at rocket firing temperatures).Although this invention is not currently intended to be limited by anytheory, it is believed that the enhanced across-ply tensile strength ofthe inventive ablative material is attributable to the orientation offibers being offset relative to the yarn axis. As a result, the ends offibers forming the yarn can entangle with the fibers of an adjacentlayer of yarn, thereby increasing the shear strength between the layersof yarn.

This invention will now be described with reference to the followingexamples, which are neither exhaustive nor exclusive of the scope ofthis invention.

EXAMPLES Example 1

3.0 dpf LYOCELL staple fibers having an average length of about 51 mmwere spun into yarns having an average denier of about 825 using acotton ring spinning machine.

Example 2

3.0 dpf LYOCELL staple fibers having an average length of about 100 mmwere spun into a yarn having denier of about 825 a using a worsted woolspinning machine.

For each of Examples 1-2, the yarns were spun into heavy tow yarns eachhaving a denier of about 825. This was accomplished by making a Ne(Number English) 6.4 spun yarn. By two-plying (twisting) the yarn into aNe 3.2 yarn, a denier of 1650 was obtained. The resulting yarns werethen woven into fabric in a square woven having a 5 harness satinconfiguration. The fabrics were then carbonized using the standardcarbonization schedules used for ablative carbon fabric filamentcellulosic fibers.

The carbonized fabric was impregnated with a phenolic resin, and inparticular phenol formaldehyde resole resin. The prepreg material was31.0-36.0 wt % phenolic resin, 13.0-17.5 wt % carbon black filler, and46.5 to 56.0 wt % carbon fabric.

The following table lists the thermal and mechanical properties ofvarious yarns, carbon cloths, and carbon cloth phenolic ablativematerials tested to compare staple cellulosic precursors with currentfilament rayon precursors.

TABLE 1 Properties of Filament and Solvent-Spun Staple Cellulosic Fibersin Carbon Cloth Phenolic Ablatives Staple Filament LYOCELL Units Rayon(Example 1) Yarn Properties Yarn Denier g/9 KM 1650 1650 Yarn Plies Ply1 2 Fibers per Yarn — 720 550 Denier per Dpf 2.3 3.0 Filament WovenFabric Properties Fabric Width cm 152 152 (inches)  (60)  (60) AreaWeight g/m² 576 576 (oz/yd²) (17.0) (17.0) Weave Pattern — 8 harnesssatin 5 harness satin Carbon Fabric Properties Fabric Width cm 109 117(inches)  (43)  (46) Area Weight g/m² 271 271 (oz/yd²)  (8.0)  (8.0)Carbon Content % 97.7 95.7 Prepreg Properties Carbon Content % 50.6 45.6Resin Content % 34.2 38.4 Filler Content % 15.2 17.0 Cured Composite andAblative Properties Across Ply MPa 26.5 32.2 Tensile (@ (psi) (3837) (4665)  21° C. or 70° F.) Across Ply MPa 2.12 5.21 Tensile (@ (psi)(307) (756) 399° C. or 750° F.) Interlaminar MPa 39.7 50.1 ShearStrength (psi) (5760)  (7267)  Nozzle Erosion μm/sec 171 173 Rate*(mils/s) (6.74) (6.82) Total Heat mm 14.1 13.1 Effected Depth* (inches)(0.556) (0.516) *Based upon solid fuel rocket test motor firing of 35seconds at 900 psi.

As shown in the above Table, the carbon cloth phenolic ablatives formedfrom staple LYOCELL fibers in accordance with this present inventionexhibited much higher across ply tensile strengths than the conventionalcontinuous filament viscose rayon precursor at room temperature of 21°C. and operating temperatures of 399° C.

The foregoing detailed description of the preferred embodiments of theinvention has been provided for the purposes of illustration anddescription. It is not intended to be exhaustive or exclusive in itsdescription of the precise embodiments disclosed. The embodiments werechosen and described in order to best explain the principles of theinvention and its practical application, thereby enabling others skilledin the art to understand the invention for various embodiments and withvarious modifications covered within the spirit and scope of theappended claims.

We claim:
 1. A method for insulating or thermally protecting a rocketmotor assembly comprising a rocket motor casing, a solid propellant, anda nozzle assembly, said process comprising (a) forming a rocket motorablative material from a prepreg comprising at least one impregnatingresin matrix and, as a precursor prior to carbonization, either cardedand yarn spun solvent-spun staple cellulosic fibers, solvent-spuncellulosic filaments, or a combination thereof and (b) insulating orlining a portion of the rocket motor assembly with the rocket motorablative material.
 2. The process of claim 1, wherein the ablativematerial comprises the solvent-spun staple cellulosic fibers.
 3. Theprocess of claim 2, the solvent-spun staple cellulosic fibers have anaverage length in a range of from 38 mm to 225 mm, and are spun intoyarn having a denier per fiber in a range of from 1.1 dpf to 6.0 dpf. 4.The process of claim 1, wherein the precursor is made by solventspinning with N-methylmorpholene-N-oxide.
 5. The process of claim 1,wherein the solvent-spun staple cellulosic fibers are untreated.
 6. Theprocess of claim 1, wherein said insulating or lining step comprisesapplying the rocket motor ablative material between the solid propellantand the casing surrounding the solid propellant.
 7. The process of claim1, wherein said insulating or lining step comprises applying theablative material as a bulk ablative material of an exit nozzle liner.8. The process of claim 1, wherein said insulating or lining stepcomprising applying the ablative material as a bulk ablative material ofa re-entry vehicle nose cone.
 9. The process of claim 1, furthercomprising carbonizing the prepreg at at least 1350° C.
 10. The processof claim 1, wherein the prepreg comprises 31.0-36.0 wt % phenolic resin,13.0-17.5 wt % carbon black filler, and 46.5 to 56.0 wt % carbon fabric.11. A rocket motor ablative material comprising a prepreg, said prepregcomprising at least one carbon-based reinforcement structure impregnatedwith at least one resin, said reinforcement structure being formed from,as a precursor prior to carbonization, either carded and yarn-spunsolvent-spun staple cellulosic fibers, solvent-spun cellulosicfilaments, or a combination thereof.
 12. The rocket motor ablativematerial of claim 11, wherein the ablative material comprises thesolvent-spun staple cellulosic fibers.
 13. The rocket motor ablativematerial of claim 12, wherein the solvent-spun staple cellulosic fibershave an average length in a range of from 51 to 225 mm, and are spuninto yarn having a denier per filament in a range of from 1.1 dpf to 6.0dpf.
 14. The rocket motor ablative material of claim 11, wherein theprecursor is made by solvent spinning with N-methylmorpholene-N-oxide.15. The rocket motor ablative material of claim 11, wherein thesolvent-spun staple cellulosic fibers are untreated.
 16. A rocket motorassembly comprising the ablative material of claim
 11. 17. The rocketmotor assembly of claim 16, wherein the ablative material is constructedand arranged as a bulk ablative material of an exit nozzle liner. 18.The rocket motor assembly of claim 16, wherein the ablative material isconstructed and arranged as a bulk ablative material of a re-entryvehicle nose cone.
 19. The rocket motor assembly of claim 16, whereinthe ablative material is interposed between a solid propellant andcasing of the rocket motor assembly.